An axial compressor for a domestic appliance can be designed to be smaller than an equivalent centrifugal compressor. However, the performance of such a compact axial compression system is limited by increased viscous losses and reduced flow turning at low Reynolds numbers ($Re$). In domestic appliance compressors, $Re$ is typically in the range $10^4$ - $10^5$. Although the aerodynamics of isolated aerofoils operating at these $Re$ have been studied extensively, the flow fields within low $Re$ axial compressors have not been investigated in detail. This dissertation aims to develop an improved understanding of loss variation at low $Re$ and to explore how the losses can be reduced through design changes. Experiments on a 5 times scaled-up single stage axial compressor have been conducted across a range of $Re$ of $10^4$ - $10^5$. The flow field has been characterised using detailed area traverses with a miniaturised five-hole probe at the rotor inlet, rotor exit and stator exit and a miniature hot-wire at the rotor exit. The probe was specifically designed and calibrated for the scale of the experiments and methods to improve the accuracy of the measurements have been applied including a probe geometry correction. The traverse experiments were performed at the design operating condition ($\phi=0.55$ and $Re= 6\times10^4$) and at a condition close to stall for a datum stage design, a stage with an improved stator design and two stators with compound lean. It was found that losses in the rotor were greater than the stator losses across the whole range of $Re$. As expected, the loss decreased with increasing $Re$ for both the stator and rotor. The losses were also increased by three-dimensional flow, with typical loss coefficients at the hub and tip of the blade rows in the range of $20-30\%$. A major contributor to the rotor loss was an unexpected hub separation that increased in size as $Re$ was reduced. At higher $Re$, the major loss sources were the rotor tip leakage, the stator wake and the stator hub separation. The results indicate that an improved stator design that accounts for the actual, measured, rotor exit flow field at low $Re$ could reduce the $Re$ at which blade row losses start to rise dramatically as well as reduce the loss across all $Re$. The improved stator design was better matched to the radial distribution of rotor exit flow angle, which led to a decrease in stator loss across all $Re$. For all stator designs, however, the measured stage stall margin was identical at all $Re$. This, along with the increase in velocity deficit in the rotor tip region at off-design indicates that stall occurred in the rotor and was neither $Re$ nor stator design dependent. The introduction of compound lean to the the stator design had the expected result of decreasing the endwall corner separation loss and increasing midspan losses. The experiments have shown that there was a loss increase in both the midspan and casing region much greater than the corresponding decrease in the stator hub. Also the mass flow redistribution in the experiments was larger that the redistribution predicted by the CFD. Three-dimensional RANS computations at low $Re$ of the same designs as experimentally studied were also conducted in order to investigate the predictive accuracy of industry standard CFD. The simulation results predicted the overall loss distribution but overestimated the end-wall losses and failed to capture the drop in stage performance at low $Re$. The differences with the experiments were caused by the inherent limitations of a fully turbulent solver that cannot reproduce transitional flow-features. Similarly to the experiments, there was no stall margin dependency on $Re$ in the simulations. This thesis has shown that with axial compressors designed specifically for low $Re$, the $Re$ at which the losses start increasing exponentially can be shifted from $10\times10^4$ to $ 4\times10^4$. The loss increase is predominantly caused by the rotor hub corner separation.

A direct numerical simulation study has been conducted to examine the flow that develops on long circular cylinders that are aligned, or nearly aligned, with the freestream. Results are presented for turbulent boundary layers and vortex - shedding yawed flow. Although flows of these types occur in a range of engineering applications, they remain relatively unexplored compared with flat - plate flow. The numerical scheme employed for solution of the governing Navier - Stokes equations is similar to that used in some previously published simulations, but here rather different boundary conditions are adopted. At the outer edge of the cylindrical computational domain, the imposed boundary conditions confine the vorticity field within a finite radius while allowing the continuous velocity field to converge to the free - stream velocity at large distances from the cylinder. Axial flows are considered with radius Reynolds numbers in the range 311 to 20800, ratios of boundary layer thickness to cylinder radius in the range 0.15 to 27.5, and boundary layer thicknesses of between 160 and 800 viscous units ( v / u [subscript τ] ). The mean - flow and turbulence statistics for axisymmetric boundary layers are found to differ significantly from flat - plate results when the boundary layer is strongly curved, that is when the boundary layer is thick in relation to the cylinder radius. The effects of curvature are mainly observed in the outer flow except when the cylinder radius in viscous units is small. Particular attention is given to the assessment of similarity scaling relations for the mean velocity profile, velocity fluctuation statistics and temporal wall - pressure spectra. Structural features of axisymmetric turbulence are examined by inspection of instantaneous flow fields, correlation functions and conditionally - averaged flow structures. In very thick boundary layers on thin cylinders, the simulations reveal evidence of large - scale fluid motion across the cylinder, although the mechanisms of turbulence generation do not appear to be significantly different from those in flatplate flow. Simulations of turbulence in near - axial flow over cylinders are considered with radius Reynolds numbers up to 674 and yaw angles up to 0.5 degrees. No previous flow simulations of this kind are reported in the literature. The mean - flow and turbulence statistics are found to depart rapidly from axisymmetry as the yaw angle increases. The quality of the calculated results suggests that the computational procedure is suitable for use in a more comprehensive investigation of near - axial flow over cylinders. For cylinders inclined at sufficiently large yaw angles to the free - stream, turbulent boundary layer flow gives way to oblique vortex - shedding from the cylinder. Simulated flow fields corresponding to a radius Reynolds number of 311 and a yaw angle of 3 degrees are examined to reveal the three - dimensional structure of the flow. The results suggest that the oscillating flow fields in the cylinder wake have the character of a wave travelling in the axial direction at the same speed as the axial component of the free - stream. / Thesis (Ph.D.)--School of Mechanical Engineering, 2006.

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Thesis (M.S. in Mechanical Engineering)--University of Michigan, 2001. / Includes bibliographical references (p. 49). Also available online.

Steady, three-dimensional viscous numerical analysis of airflow over a rectangular NACA 0012 base wing (BW) with a rounded tip and with three NACA 0015 tip-sails (WTS) is performed. The flow physics and aerodynamic forces are studied at Reynolds numbers (Re) of 60,000 and 600,000, angles of attack (α) of 0, 5, 7.5, and 10°, and two sets of tip-sail dihedral angles (leading to trailing tip-sail): 50, 45, and 40° and 60, 45, and 30°. The Shear Stress Transport turbulence and intermittency-transition Reynolds number transitional turbulence models were used. For α > 0°, the WTS produced higher lift coefficients (CL) and drag coefficients (CD) than the BW. At Re = 600,000 and α > 0°, the CL/CD was higher for the WTS than the BW. Good agreement was seen with experimental data at Re = 600,000 for the BW results and the WTS CL but not the WTS CD. / October 2015

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